The present invention relates to stationary rings surrounding gas passages in a gas turbine, and more particularly it relates to cooling stationary rings in a gas turbine.
A gas turbine, in particular a high pressure turbine of a turbomachine, typically comprises a plurality of stationary vanes alternating with a plurality of moving blades in the passage for hot gas coming from the combustion chamber of the turbomachine. The moving blades on the turbine are surrounded over their entire circumference by a stationary ring that is generally made up of a plurality of ring segments. These ring segments define part of the flow passage for hot gas passing through the blades of the turbine.
The ring segments of the turbine are thus subjected to the high temperatures of the hot gas coming from the combustion chamber of the turbomachine. To enable the turbine ring to withstand the temperature and mechanical stresses to which it is subjected, it is necessary to provide the ring segments with cooling devices.
One of the known methods of cooling consists in feeding cooling air to an impact plate mounted on the bodies of the ring segments. The plate is provided with a plurality of orifices for passing air which, under the pressure difference between the sides of the plate, comes to cool the ring segment by impact. The cooling air is then exhausted into the hot gas passage via holes formed through the ring segment.
Such a method does not enable effective and uniform cooling of the ring segments to be obtained, particularly at the upstream ends of the ring segments which constitute a zone that is particularly exposed to hot gas. This therefore has an affect on the lifetime of the ring segment. Furthermore, that technology requires too great an amount of cooling air to be taken, thereby decreasing the performance of the turbine.